This invention generally relates to a gas turbine engine, and more particular to a nacelle assembly for a turbofan gas turbine engine.
In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages which extract energy from the hot combustion gases. A fan supplies air to the compressor.
Combustion gases are discharged from the turbofan engine through a core exhaust nozzle and a quantity of fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle assembly surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust is provided from the combustion gases discharged through the core exhaust nozzle.
The fan section of a turbofan gas turbine engine may be geared to control a tip speed of the fan section. The ability to reduce the fan section tip speed results in decreased noise due to the fan section tip speed being lower than the speed of the rotating compressor. Controlling the fan section tip speed allows the fan section to be designed with a larger diameter, which further decreases noise. However, the nacelle assembly of the turbofan engine must be large enough to support the large diameter fan section.
It is known in the field of aircraft gas turbine engines that the performance of a turbofan engine varies during diversified conditions experienced by the aircraft. An inlet lip section located at the foremost end of the turbofan nacelle assembly is typically designed to enable operation of the turbofan engine and reduce the separation of airflow from the inlet lip section of the nacelle assembly during these diversified conditions. For example, the inlet lip section requires a “thick” inlet lip section to support operation of the engine during specific flight conditions, such as crosswind conditions, take-off and the like. Disadvantageously, the “thick” inlet lip section may reduce the efficiency of the turbofan engine during normal cruise conditions of the aircraft. As a result, the maximum diameter of the nacelle assembly may be approximately 10-20% larger than needed at cruise conditions.
In addition, boundary layer separation is a common problem associated with “thin” inlet lip sections. Boundary layer separation occurs where airflow communicated through the inlet lip section separates from the outer and/or inner flow surfaces of the inlet lip section, which may cause engine stall, the loss of the capability to generate thrust, and may decrease engine efficiency.
Attempts have been made to reduce the onset of boundary layer separation within the nacelle assembly. For example, small vortex generators are known which increase the velocity gradient of oncoming airflow near the effective boundary layer of the inlet lip section. In addition, synthetic jets are known which introduce an airflow at the boundary layer to increase the velocity gradient of the oncoming airflow near the boundary separation point. However, these attempts have proved complex, expensive and have not fully reduced the onset of boundary layer separation.
Accordingly, it is desirable to improve the performance of a turbofan gas turbine engine during diversified conditions to provide a nacelle assembly having a reduced thickness, reduced weight and reduced drag.